The present invention relates generally to gas turbine engines, and, more specifically, to high pressure turbine nozzles.
In a typical turbofan aircraft gas turbine engine, air is pressurized in a multistage axial compressor, mixed with fuel in a combustor, and ignited for generating hot combustion gases which flow downstream through a high pressure (HP) turbine nozzle which turns and accelerates the gases for energy extraction therefrom in downstream high pressure turbine rotor blades. The gases then flow through a low pressure turbine which extracts additional energy for powering a fan to produce propulsion thrust for powering the aircraft in flight. The many components disposed in the flowpath of the hot combustion gases are heated thereby and must be suitably protect therefrom.
For example, thermal barrier coating (TBC) is a ceramic material having various conventional compositions which may be applied in thin layers atop the various components for providing thermal protection thereof. The TBC may be conventionally applied using plasma spray techniques or physical vapor deposition.
The TBC provides a barrier between the hot combustion gases and the underlying metal of the specific components and provides thermal insulation for reducing the maximum temperature experienced by the component for improving the useful life thereof in the engine.
Since the TBC is a ceramic material it is also relatively brittle compared to the underlying metal substrate, and therefore, its integrity and corresponding durability is in large part determined by the strength and operating experience of the underlying component. For example, the HP turbine nozzle vanes receive the hottest temperature combustion gases from the combustor and require corresponding protection.
Various configurations of turbine nozzle vanes have enjoyed many years of successful commercial use when protected with TBC. Typical nozzle vanes are radially straight and twist relative to the trailing edges thereof for defining converging channels therebetween ending in throats of minimum flow area through which the combustion gases are turned and accelerated toward the turbine rotor blades.
The TBC may be applied along the suction sides of the vanes as well as along the pressure sides exclusive of the vane throat in conventional practice. The nozzle throat area is a critical design parameter which affects the operating efficiency of the turbine and therefore the entire engine. The individual vane throat areas arid the collective throat area must be maintained within a suitable narrow tolerance for optimum engine efficiency. Since TBC is conventionally applied with a thickness tolerance of plus or minus a few mils, this tolerance variation would be unacceptable in maintaining consistent nozzle throat area, and therefore the TBC is not provided on the suction sides of the vanes near the leading edges which forms one boundary of the vane throat, with the other boundary being defined by the pressure side along the trailing edge of the next adjacent nozzle vane.
In a recent development enjoying successful commercial use in this country for several years, a 3-D nozzle vane includes a trailing edge having a bow instead of being straight for increasing total pressure and momentum in the combustion gases at the root of the vanes near their supporting inner bands. The 3-D vane twists about its leading edge between the inner and outer bands and also leans along the trailing edge to define the bow. Three dimensional computer analysis software is available for defining the specific curvature and extent of the bow to increase gas flow momentum near the inner band for improving the overall efficiency of the turbine and engine.
In order to protect the 3-D bowed vanes against the high temperatures of the combustion gases, the vanes have included full coverage TBC along both their pressure and suction sides exclusive of the vane throats. The application of the TBC to the bowed nozzle vane is even more critical than for straight vanes since differential temperatures commonly occurring over the surfaces of the vane can create corresponding thermal stress and distortion therein. The bowed trailing edge, for example, is now subject to bending loads due to its non-straight configuration, and is therefore also subject to distortion in its curvature. Since the trailing edge defines one boundary of the vane throat, any variation in that boundary changes the throat area which can undesirably decrease the efficiency of the turbine and the engine.
Throat area changes also alter the total pressure drop across the turbine nozzle and correspondingly increase loads in the thrust bearing which reacts the differential loads between the compressor and the turbine rotor.
Several years of commercial experience of the full coverage TBC 3-D bowed turbine nozzle has shown failure in the TBC such as premature spallation along the leading edges of the vanes.
Accordingly, it is desired to eliminate the premature failure of the TBC in the 3-D bowed nozzle vane without adversely affecting aerodynamic performance or efficiency of the nozzle and engine, and obtaining a suitable useful life of the turbine nozzle.
A turbine nozzle includes outer and inner bands between which extend a plurality of vanes for channeling combustion gases. Each of the vanes includes leading and trailing edges, and pressure and suction sides extending therebetween, and also a bow along the trailing edge to increase pressure in the gases adjacent the inner band. The vanes also include a thermal barrier coating (TBC) selectively disposed solely along the suction side between the leading and trailing edges.